Ceramic Matrix Composite Airfoil Cooling

ABSTRACT

Ceramic matrix composite airfoils for gas turbine engines are provided. In an exemplary embodiment, an airfoil includes opposite pressure and suction sides extending radially along a span. The pressure and suction sides define an outer surface of the airfoil. The airfoil further includes opposite leading and trailing edges extending radially along the span, the pressure and suction sides extending axially between the leading and trailing edges. The airfoil also includes a filler pack defining the trailing edge; the filler pack comprises a ceramic matrix composite material. Moreover, the airfoil includes a plenum defined within the airfoil for receiving a flow of cooling fluid, and a cooling passage defined within the filler pack for directing the flow of cooling fluid from the plenum to the outer surface of the airfoil. Methods for forming airfoils for gas turbine engines also are provided.

FIELD OF THE INVENTION

The present subject matter relates generally to a gas turbine engine, ormore particularly to features for cooling internal components of gasturbine engines. Most particularly, the present subject matter relatesto trailing edge cooling for gas turbine engine airfoils.

BACKGROUND OF THE INVENTION

A gas turbine engine generally includes a fan and a core arranged inflow communication with one another. Additionally, the core of the gasturbine engine generally includes, in serial flow order, a compressorsection, a combustion section, a turbine section, and an exhaustsection. In operation, air is provided from the fan to an inlet of thecompressor section where one or more axial compressors progressivelycompress the air until it reaches the combustion section. Fuel is mixedwith the compressed air and burned within the combustion section toprovide combustion gases. The combustion gases are routed from thecombustion section to the turbine section. The flow of combustion gasesthrough the turbine section drives the turbine section and is thenrouted through the exhaust section, e.g., to atmosphere.

In general, turbine performance and efficiency may be improved byincreased combustion gas temperatures. However, increased combustiontemperatures can negatively impact the gas turbine engine components,for example, by increasing the likelihood of material failures. Thus,while increased combustion temperatures can be beneficial to turbineperformance, some components of the gas turbine engine may requirecooling features or reduced exposure to the combustion gases to decreasethe negative impacts of the increased temperatures on the components.

Film cooling gas turbine engine components, e.g., by directing a flow ofcooler fluid over the surface of the component, can help reduce thenegative impacts of elevated combustion temperatures. For example,cooling apertures may be provided throughout a component that allow aflow of cooling fluid from within the component to be directed over theouter surface of the component. However, multiple rows of cooling holesoften are required to achieve beneficial film cooling, and the multiplerows of cooling holes can be detrimental to the component structure aswell as engine performance. Also, typical drilling processes fordefining the cooling holes require increased component thicknesses toaccommodate tolerances in drill hole placement, thereby increasing theweight of and material required to produce the component. Further, knowncooling hole configurations often have only a single solution formetering the flow of cooling fluid.

Therefore, improved cooling features for gas turbine components thatovercome one or more disadvantages of existing cooling features would bedesirable. In particular, an airfoil for a gas turbine engine havingtrailing edge cooling features that minimize a thickness of a trailingedge portion of the airfoil would be beneficial. Moreover, an airfoilfor a gas turbine engine having trailing edge cooling features thatreduce cooling flow would be desirable. Further, an airfoil havingtrailing edge cooling features that minimize or reduce manufacturingtime and cost would be advantageous. Additionally, a method for formingan airfoil for a gas turbine engine where the airfoil has features forimproved trailing edge cooling would be useful.

BRIEF DESCRIPTION OF THE INVENTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

In one exemplary embodiment of the present disclosure, a ceramic matrixcomposite airfoil for a gas turbine engine is provided. The airfoilincludes opposite pressure and suction sides extending radially along aspan. The pressure and suction sides define an outer surface of theairfoil. The airfoil further includes opposite leading and trailingedges extending radially along the span. The pressure and suction sidesextend axially between the leading and trailing edges. The airfoil alsoincludes a filler pack defining the trailing edge; the filler packcomprises a ceramic matrix composite material. Moreover, the airfoilincludes a plenum defined within the airfoil for receiving a flow ofcooling fluid, and a cooling passage defined within the filler pack fordirecting the flow of cooling fluid from the plenum to the outer surfaceof the airfoil.

In another exemplary embodiment of the present disclosure, a ceramicmatrix composite airfoil for a gas turbine engine is provided. Theairfoil includes opposite pressure and suction sides extending radiallyalong a span; the pressure and suction sides defining an outer surfaceof the airfoil. The airfoil further includes opposite leading andtrailing edges extending radially along the span. The pressure andsuction sides extending axially between the leading and trailing edges.Also, the airfoil includes a filler pack defining the trailing edge; thefiller pack comprises a ceramic matrix composite material. Moreover, theairfoil includes a plenum defined within the airfoil for receiving aflow of cooling fluid; a chamber defined in the filler pack; a crossoveraperture defined from the plenum to the chamber; and an ejectionaperture defined in the filler pack from the chamber to the outersurface of the airfoil adjacent the trailing edge. The crossoveraperture facilitates the flow of cooling fluid from the plenum to thechamber, and the ejection aperture is in fluid communication with thechamber to direct the flow of cooling fluid from the plenum to the outersurface of the airfoil.

In a further exemplary embodiment of the present disclosure, a methodfor forming an airfoil for a gas turbine engine is provided. The methodincludes laying up a ceramic matrix composite material to form anairfoil preform assembly. The airfoil preform assembly includes oppositepressure and suction sides extending radially along a span, oppositeleading and trailing edges extending radially along the span, and aplenum defined within the airfoil preform assembly. The pressure andsuction sides extend axially between the leading and trailing edges. Themethod further includes processing the airfoil preform assembly toproduce the airfoil. A cooling passage is defined within the airfoil.The cooling passage is defined from the plenum to the trailing edge ofthe airfoil.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 provides a schematic cross-sectional view of an exemplary gasturbine engine according to various embodiments of the present subjectmatter.

FIG. 2 provides a side perspective view of a turbine rotor bladeaccording to an exemplary embodiment of the present subject matter.

FIG. 3 provides a perspective view of a turbine nozzle segment accordingto an exemplary embodiment of the present subject matter.

FIG. 4 provides a cross-section view of a portion of an airfoil of theturbine nozzle segment, taken along the line 4-4 of FIG. 3, according toan exemplary embodiment of the present subject matter.

FIG. 5 provides a cross-section view of a portion of the airfoil of theturbine nozzle segment, taken along the line 5-5 of FIG. 3, according toan exemplary embodiment of the present subject matter.

FIG. 6 provides a cross-section view of a portion of the airfoil of theturbine nozzle segment, taken along the line 6-6 of FIG. 5, according toan exemplary embodiment of the present subject matter.

FIG. 7 provides the cross-section view of the portion of the airfoil ofthe turbine nozzle segment of FIG. 6 according to another exemplaryembodiment of the present subject matter.

FIG. 8 provides a chart illustrating a method for forming an airfoil ofa gas turbine engine according to an exemplary embodiment of the presentsubject matter.

FIG. 9 provides a chart illustrating a portion of the method of FIG. 8according to an exemplary embodiment of the present subject matter.

FIG. 10 provides a cross-section view of an airfoil preform assemblyaccording to an exemplary embodiment of the present subject matter.

DETAILED DESCRIPTION OF THE INVENTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention. As used herein, theterms “first,” “second,” and “third” may be used interchangeably todistinguish one component from another and are not intended to signifylocation or importance of the individual components. The terms“upstream” and “downstream” refer to the relative direction with respectto fluid flow in a fluid pathway. For example, “upstream” refers to thedirection from which the fluid flows and “downstream” refers to thedirection to which the fluid flows.

Referring now to the drawings, wherein identical numerals indicate thesame elements throughout the figures, FIG. 1 is a schematiccross-sectional view of a gas turbine engine in accordance with anexemplary embodiment of the present disclosure. More particularly, forthe embodiment of FIG. 1, the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “turbofan engine 10.” Asshown in FIG. 1, the turbofan engine 10 defines an axial direction A(extending parallel to a longitudinal centerline 12 provided forreference) and a radial direction R. In general, the turbofan 10includes a fan section 14 and a core turbine engine 16 disposeddownstream from the fan section 14.

The exemplary core turbine engine 16 depicted generally includes asubstantially tubular outer casing 18 that defines an annular inlet 20.The outer casing 18 encases, in serial flow relationship, a compressorsection including a booster or low pressure (LP) compressor 22 and ahigh pressure (HP) compressor 24; a combustion section 26; a turbinesection including a high pressure (HP) turbine 28 and a low pressure(LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure(HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HPcompressor 24. A low pressure (LP) shaft or spool 36 drivingly connectsthe LP turbine 30 to the LP compressor 22.

For the depicted embodiment, fan section 14 includes a variable pitchfan 38 having a plurality of fan blades 40 coupled to a disk 42 in aspaced apart manner. As depicted, fan blades 40 extend outward from disk42 generally along the radial direction R. Each fan blade 40 isrotatable relative to disk 42 about a pitch axis P by virtue of the fanblades 40 being operatively coupled to a suitable actuation member 44configured to vary the pitch of the fan blades 40. Fan blades 40, disk42, and actuation member 44 are together rotatable about thelongitudinal axis 12 by LP shaft 36 across a power gear box 46. Thepower gear box 46 includes a plurality of gears for stepping down therotational speed of the LP shaft 36 to a more efficient rotational fanspeed.

Referring still to the exemplary embodiment of FIG. 1, disk 42 iscovered by rotatable front nacelle 48 aerodynamically contoured topromote an airflow through the plurality of fan blades 40. Additionally,the exemplary fan section 14 includes an annular fan casing or outernacelle 50 that circumferentially surrounds the fan 38 and/or at least aportion of the core turbine engine 16. It should be appreciated thatnacelle 50 may be configured to be supported relative to the coreturbine engine 16 by a plurality of circumferentially-spaced outletguide vanes 52. Moreover, a downstream section 54 of the nacelle 50 mayextend over an outer portion of the core turbine engine 16 so as todefine a bypass airflow passage 56 therebetween.

During operation of the turbofan engine 10, a volume of air 58 entersturbofan 10 through an associated inlet 60 of the nacelle 50 and/or fansection 14. As the volume of air 58 passes across fan blades 40, a firstportion of the air 58 as indicated by arrows 62 is directed or routedinto the bypass airflow passage 56 and a second portion of the air 58 asindicated by arrows 64 is directed or routed into the LP compressor 22.The ratio between the first portion of air 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the secondportion of air 64 is then increased as it is routed through the highpressure (HP) compressor 24 and into the combustion section 26, where itis mixed with fuel and burned to provide combustion gases 66.

The combustion gases 66 are routed through the HP turbine 28 where aportion of thermal and/or kinetic energy from the combustion gases 66 isextracted via sequential stages of HP turbine stator vanes 68 that arecoupled to the outer casing 18 and HP turbine rotor blades 70 that arecoupled to the HP shaft or spool 34, thus causing the HP shaft or spool34 to rotate, thereby supporting operation of the HP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where asecond portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LP turbine stator vanes 72that are coupled to the outer casing 18 and LP turbine rotor blades 74that are coupled to the LP shaft or spool 36, thus causing the LP shaftor spool 36 to rotate, thereby supporting operation of the LP compressor22 and/or rotation of the fan 38.

The combustion gases 66 are subsequently routed through the jet exhaustnozzle section 32 of the core turbine engine 16 to provide propulsivethrust. Simultaneously, the pressure of the first portion of air 62 issubstantially increased as the first portion of air 62 is routed throughthe bypass airflow passage 56 before it is exhausted from a fan nozzleexhaust section 76 of the turbofan 10, also providing propulsive thrust.The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section32 at least partially define a hot gas path 78 for routing thecombustion gases 66 through the core turbine engine 16.

In some embodiments, components of turbofan engine 10, particularlycomponents within hot gas path 78, may comprise a ceramic matrixcomposite (CMC) material, which is a non-metallic material having hightemperature capability. Exemplary CMC materials utilized for suchcomponents may include silicon carbide, silicon, silica, or aluminamatrix materials and combinations thereof. Ceramic fibers may beembedded within the matrix, such as oxidation stable reinforcing fibersincluding monofilaments like sapphire and silicon carbide (e.g.,Textron's SCS-6), as well as rovings and yarn including silicon carbide(e.g., Nippon Carbon's NICALON®, Ube Industries' TYRANNO®, and DowCorning's SYLRAMIC®), alumina silicates (e.g., Nextel's 440 and 480),and chopped whiskers and fibers (e.g., Nextel's 440 and SAFFIL®), andoptionally ceramic particles (e.g., oxides of Si, Al, Zr, Y, andcombinations thereof) and inorganic fillers (e.g., pyrophyllite,wollastonite, mica, talc, kyanite, and montmorillonite). As furtherexamples, the CMC materials may also include silicon carbide (SiC) orcarbon fiber cloth.

CMC materials may be used for various components of the engine, forexample, airfoils in the turbine, compressor, and/or fan regions. Thecompressor and turbine generally include rows of airfoils that arestacked axially in stages. Each stage includes a row ofcircumferentially spaced stator vanes and a rotor assembly that rotatesabout centerline 12 of engine 10. Turbine nozzles, comprising statorvanes extending between inner and outer bands, direct the hot combustiongas in a manner to maximize extraction at the adjacent downstreamturbine blades. In various embodiments of engine 10, the nozzles and/orturbine blades, including their associated airfoils, may be CMCcomponents. Of course, other components of turbine engine 10 also may beformed from CMC materials.

Referring now to FIG. 2, a side, perspective view of a turbine rotorblade 74, a portion of a turbine rotor assembly, is provided accordingto an exemplary embodiment of the present subject matter. As previouslydescribed, LP turbine 30 includes sequential stages of turbine statorvanes 72 coupled to outer casing 18 and turbine rotor blades 74 coupledto shaft or spool 36. Each blade 74 includes an airfoil 80 having aconcave pressure side 82 opposite a convex suction side 84. Oppositepressure and suction sides 82, 84 of each airfoil 80 extend radiallyalong a blade span S from a root 86 to a tip 87 and define an outersurface 85 of airfoil 80. As depicted, root 86 is the radially innermostportion of airfoil 80 and tip 87 is the radially outermost portion ofairfoil 80. Moreover, as further shown in FIG. 2, pressure and suctionsides 82, 84 of airfoil 80 extend axially between a leading edge 88 andan opposite trailing edge 90. Leading and trailing edges 88, 90 extendradially from root 86 to tip 87. Further, leading edge 88 defines aforward end of airfoil 80 (labeled Fwd in the Figures), and trailingedge 90 defines an aft end of airfoil 80 (labeled Aft in the Figures).Further, airfoil 80 defines a chord C extending axially between oppositeleading and trailing edges 88, 90. Moreover, airfoil 80 defines a widthW between pressure side 82 and suction side 84. The width W of airfoil80 may vary along the span S.

Each blade 74 is coupled to shaft or spool 36 via root 86. Moreparticularly, root 86 is coupled to a turbine rotor disk (not shown),which in turn is coupled to shaft or spool 36 (FIG. 1). It will bereadily understood that, as is depicted in FIG. 2 and is generallywell-known in the art, root 86 may define a projection 89 having adovetail or other shape for receipt in a complementarily shaped slot inthe turbine rotor disk to couple blade 74 to the disk. Of course, eachblade 74 may be coupled to the turbine rotor disk and/or shaft or spool36 in other ways as well. In any event, blades 74 are coupled to theturbine rotor disks such that a row of circumferentially adjacent blades74 extends radially outward from the perimeter of each disk, i.e.,adjacent blades 74 within a blade row are spaced apart from one anotheralong a circumferential direction M and each blade 74 extends from thedisk along the radial direction R. As such, the turbine rotor disk andouter casing 18 form an inner end wall and an outer end wall,respectively, of hot gas path 78 through the turbine assembly.

Referring now to FIG. 3, a perspective view is provided of a turbinenozzle segment. A turbine stator is formed by a plurality of turbinenozzle segments that are abutted at circumferential ends to form acomplete ring about centerline 12. Each nozzle segment may comprise oneor more vanes, such as vanes 68 of HP turbine 28 or vanes 72 of LPturbine 30, that extend between an outer band and an inner band aspreviously described. FIG. 3 depicts an exemplary turbine nozzle segment67 of HP turbine 28. Nozzle segment 67 includes outer band 67 a andinner band 67 b, between which extends stator vanes 68. Each stator vane68 includes an airfoil 80, which has the same features as airfoil 80described above with respect to blade 74. For example, airfoil 80 ofvane 68 has a pressure side 82 opposite a suction side 84. Oppositepressure and suction sides 82, 84 of each airfoil 80 extend radiallyalong a span from a vane root at inner band 67 b to a vane tip at outerband 67 a. Moreover, pressure and suction sides 82, 84 of airfoil 80extend axially between a leading edge 88 and an opposite trailing edge90. Airfoil 80 further defines a chord extending axially betweenopposite leading and trailing edges 88, 90. Moreover, airfoil 80 definesa width between pressure side 82 and suction side 84, which may varyalong the span.

It will be appreciated that, although airfoil 80 of vane 68 may have thesame features as airfoil 80 of blade 74, airfoil 80 of vane 68 may havea different configuration than airfoil 80 of blade 74. As an example,the span of airfoil 80 of vane 68 may be larger or smaller than the spanof airfoil 80 of blade 74. As another example, the width and/or chord ofairfoil 80 of vane 68 may differ from the width and/or chord of airfoil80 of blade 74. Additionally or alternatively, airfoils 80 of LP statorvanes 72 and/or airfoils 80 of HP turbine rotor blades 70 may differ insize, shape, and/or configuration from airfoils 80 of HP stator vanes 68and LP turbine rotor blades 74. However, it also should be understoodthat, while airfoils 80 may differ in size, shape, and/or configuration,the subject matter described herein may be applied to any airfoil withinengine 10, as well as other suitable components of engine 10.

FIG. 4 provides a cross-sectional view of a portion of airfoil 80 ofstator vane 68, taken along the line 4-4 of FIG. 3. More particularly,airfoil 80 is a CMC component of engine 10. As illustrated, airfoil 80is defined by a first plurality of plies 92 and a filler pack 94.Airfoil 80 further comprises a second plurality of plies 96 defining aplenum 98 within airfoil 80. Plenum 98 receives a flow of cooling fluidF, e.g., a flow of pressurized air diverted from HP compressor 24.Further, it will be appreciated that the first plurality of plies 92also may be referred to as airfoil plies 92, and the second plurality ofplies 96 also may be referred to as plenum plies 96.

Continuing with FIG. 4, each of the plurality of airfoil plies 92extends from pressure side 82 to suction side 84 of airfoil 80. In theembodiment illustrated in FIGS. 3 and 4, each ply 92 wraps from pressureside 82 to suction side 84 around leading edge 88, or from suction side84 to pressure side 82 around leading edge 88, and thereby defineleading edge 88 of airfoil 80. However, airfoil plies 92 do not extendfrom or to trailing edge 90 on pressure side 82 or extend from or totrailing edge 90 on suction side 84. Rather, filler pack 94 definestrailing edge 90, i.e., filler pack 94 extends to trailing edge 90. Assuch, airfoil plies 92 partially define pressure and suction sides 82,84 of airfoil 80, and filler pack 94 also partially defines pressure andsuction sides 82, 84. In some embodiments, filler pack 94 may comprisetwo halves, i.e., filler pack 94 may include a first portion 94 a and asecond portion 94 b. In other embodiments, filler pack 94 may be asingle part, or in still other embodiments, filler pack 94 may comprisemore than two portions or parts.

Preferably, airfoil and plenum plies 92, 96 contain continuous CMCfibers along their lengths. Continuous fiber CMC plies can help avoidrelying on the interlaminar capability of the airfoil material to resiststresses on the airfoil. The continuous fibers may be maintained, e.g.,by wrapping each airfoil ply 92 around leading edge 88. Plenum plies 96may be wrapped around a mandrel or other appropriate support to helpmaintain continuous fibers in plies 96 as airfoil 80 is formed.

It should be appreciated that, in general, filler packs 94 may be formedfrom any suitable material and/or by using any suitable process. Forexample, in several embodiments, each filler pack 94 may be formed froma suitable fiber-reinforced composite material, such as a carbon orglass fiber-reinforced composite material. For instance, one or morefabric plies may be wrapped in a suitable manner to form one or morefiller packs 94 defining the desired shape of trailing edge 90, such asby shaping suitable ply packs to form each filler pack 94. In anotherembodiment, discontinuous materials, such as short or chopped fibers,particulates, platelets, whiskers, etc., may be dispersed throughout asuitable matrix material and used to form each filler pack 94. Fillerpack(s) 94 may have any suitable configuration for providing atransition between adjacent plies.

Additionally, it should be appreciated that, in several embodiments,each filler pack 94 may correspond to a pre-fabricated component. Insuch embodiments, the filler pack(s) may be laid up with the plies usedto define pressure and suction sides 82, 84 of airfoil 80 duringmanufacturing of the nozzle segment 67 or rotor blade 74. Alternatively,each filler pack 94 may be assembled or otherwise formed with airfoil80. For instance, when filler pack 94 is formed from one or more fabricplies, the plies may be laid up within airfoil 80 together with theplies being used to create the airfoil structure.

Referring now to FIGS. 5 and 6, FIG. 5 depicts a cross-sectional view ofa portion of airfoil 80, taken along the line 5-5 of FIG. 3, and FIG. 6provides a cross-sectional view of a trailing edge portion 91 of airfoil80, taken along the line 6-6 of FIG. 5. Trailing edge portion 91 isdefined adjacent trailing edge 90 at the aft end of airfoil 80; plenum98 is defined within airfoil 80 forward of trailing edge portion 91,i.e., closer to the forward end of airfoil 80 than the aft end. As shownin these figures, airfoil 80 defines cooling passages 100 for providinga flow of cooling fluid at trailing edge 90 of airfoil 80. A radiallyextending cavity 102 is defined in filler pack 94. Referringparticularly to FIG. 5, cavity 102 may be defined within filler pack 94at a generally central location between pressure side 82 and suctionside 84. That is, cavity 102 may be positioned essentially central to asolid volume of airfoil 80 defining the trailing edge portion 91 ofairfoil 80. Stated differently, trailing edge portion 91 is essentiallysolid except for cooling passages 100 defined therein, and cavity 102may be defined generally centrally within solid trailing edge portion91. Further, a crossover aperture 104 is defined in plenum plies 96 andfiller pack 94 such that crossover aperture 104 extends from plenum 98to cavity 102 and thereby facilitates a flow of cooling fluid fromplenum 98 to cavity 102. Further, an ejection aperture 106 is defined infiller pack 94. Ejection aperture 106 extends from cavity 102 to theouter surface 85 of airfoil 80 adjacent trailing edge 90. Ejectionaperture 106 is in fluid communication with cavity 102 to direct theflow of cooling fluid from plenum 98 to outer surface 85 andparticularly toward trailing edge 90 of airfoil 80.

The fluid flow F received within plenum 98 generally is cooler than thecombustion gases flowing against or over outer surface 85 of airfoil 80.Each cooling passage 100, extending from plenum 98 to outer surface 85via cavity 102, crossover aperture 104, and ejection aperture 106, formsa continuous pathway in fluid communication with plenum 98 to facilitateflowing cooling fluid F from plenum 98 to outer surface 85. As such, theflow of cooing fluid F over outer surface 85 and trailing edge 90 mayhelp reduce the temperatures to which outer surface 85 and trailing edge90 are exposed.

As illustrated in FIG. 6, a plurality of cooling passages 100 may beused throughout the trailing edge portion 91 of airfoil 80. Morespecifically, cavity 102 may extend radially through filler pack 94 anda plurality of ejection apertures 106 may be defined in filler pack 94from outer surface 85, e.g., at or adjacent trailing edge 90, to cavity102. Similarly, a plurality of crossover apertures 104 may be definedfrom plenum 98 to cavity 102 to provide the flow F of cooling fluid fromplenum 98 to cavity 102, which may then be ejected to outer surface 85of airfoil 80 via ejection apertures 106. As depicted in FIGS. 5 and 6,crossover apertures 104 are defined in plenum plies 96 and filler pack94. Thus, in the embodiment depicted in FIG. 6, cooling passages 100include a crossover aperture 104 defined from plenum 98 to cavity 102and an ejection aperture 106 defined from trailing edge 90 (or adjacentthereto) to cavity 102. However, an identical number of crossoverapertures 104 and ejection apertures 106 need not be provided, as shownin FIG. 6. Rather, the number of crossover apertures 104 provided inairfoil 80 may be fewer or greater than the number of ejection apertures106.

FIG. 7 provides an alternative embodiment of the trailing edge portion91 of airfoil 80 illustrated in FIG. 6. As shown in FIG. 7, rather thana cavity 102 extending radially through filler pack 94 and connecting aplurality of ejection apertures 106, a plurality of chambers 108 may beprovided such that ejection apertures 106 are not fluidly connected withone another within filler pack 94. Instead, each ejection aperture 106is in fluid communication with a chamber 108, which in turn is in fluidcommunication with a crossover aperture 104 to receive a flow F ofcooling fluid from plenum 98. Together, each chamber 108, crossoveraperture 104, and ejection aperture 106 defines a cooling passage 100.Thus, in embodiments such as illustrated in FIG. 7, each cooling passage100 defined in airfoil 80 includes a chamber 108, a crossover aperture104, and an ejections aperture 106, and airfoil 80 comprises a pluralityof cooling passages 100.

Crossover apertures 104 and ejection apertures 106 may range from about10 to about 30 mils in diameter. For example, in one embodiment, eachcrossover aperture 104 may be about 20 mils in diameter, and eachejection aperture 106 may be about 15 mils in diameter. In otherembodiments, one crossover aperture 104 may have a different diameterthan another crossover aperture 104. Alternatively or additionally, oneejection aperture 106 may have a different diameter than anotherejection aperture 106. Further, although generally described as beingsubstantially cylindrical in shape or generally circular incross-sectional shape, crossover apertures 104 and ejection apertures106, as well as cavity 102 and chambers 108, may have any appropriateshape and/or cross-section. For example, as shown in FIG. 5, cavity 102may have a generally triangular cross-sectional shape. Moreover, thenumber of each void, e.g., cavity 102, crossover aperture 104, ejectionaperture 106, and chamber 108, may vary from one airfoil to another. Asan example, airfoil 80 of turbine blade 74 may have one number ofejection apertures 106, and the airfoil of a stator vane 68 may have adifferent number of ejection apertures 106. In one example embodiment,an engine 10 may comprise airfoils having ejection apertures 106, whereejection apertures 106 of each airfoil range in number from about 10 to40 apertures, generally with larger airfoils (e.g., larger in the radialdirection R, axial direction A, circumferential direction M, or allthree directions R, A, and M) having a greater number of ejectionapertures 106.

The shape, size, and number of each void, e.g., cavity 102 and/orchamber 108, crossover aperture 104, and ejection aperture 106, may beoptimized for each airfoil. As described above, the number of ejectionapertures 106 may depend on the relative size of the airfoil. Further,the size, shape, and/or number of voids 102, 104, 106, 108 may depend onthe desired cooling effects achieved by flowing cooling fluid fromplenum 98 through voids 102, 104, 106, 108. For example, achieving highvelocity cooling fluid flow through ejection apertures 106 may increasethe heat transfer coefficient and thereby increase the rate of coolingprovided by cooling passages 100. As a result, having a larger number ofholes or voids with smaller cross-sectional areas may be beneficial.However, too many voids within the airfoil can be detrimental to thestrength of the material forming the airfoil and having too many rows ofcooling passages over airfoil 80 can increase cooling flow to an extentthat negatively impacts the performance of engine 10. Therefore, anoptimal number, shape, and size of voids 102, 104, 106, 108 providesbeneficial cooling without overly weakening the airfoil material ornegatively impacting engine performance, e.g., an optimal configurationof cooling passages 100 may decrease specific fuel consumption.

The size and/or shape of the voids forming cooling passages 100 may bedefined by various parameters of each void. For example, as shown inFIG. 5, ejection aperture 106 has a length L, which extends generallyalong the axial direction A. Also, crossover aperture 104 has a widthW_(er). Cavity 102 (or chamber 108 in embodiments having chamber 108rather than cavity 102) has a width W, adjacent crossover aperture 104,and ejection aperture 106 has a width W_(e1) adjacent cavity 102 (orchamber 108). In the depicted embodiment, width W_(ev) tapers to widthW_(e1), i.e., width W_(e1) is smaller or less than width W_(ev).Further, ejection aperture 106 has a width W_(e2) at outer surface 85,and width W_(e1) of ejection aperture 106 is smaller than width W_(e2)of ejection aperture 106, i.e., ejection aperture 106 may have a greaterwidth at outer surface 85 that tapers to a smaller width at or nearcavity 102 or chamber 108. Moreover, it will be understood that,although described as widths, the foregoing dimensions may be diametersin embodiments in which the voids are rounded or generally circular incross-sectional shape.

In addition, the size and/or shape of cavity 102 and chamber 108 may beselected to help in fabricating airfoil 80. More particularly, a largercross-sectional area of cavity 102 or chamber 108 may help in formingcrossover apertures 104 to fluidly connect plenum 98 and cavity 102 orchamber 108. For example, a cavity 102 having a larger cross-sectionalarea oriented toward a location where crossover apertures 104 will bemachined through plenum plies 96 and filler pack 94 will provide alarger target area for machining apertures 104. As shown in FIG. 5,cavity 102 may have a generally triangular cross-sectional shape, with alonger side oriented toward plenum 98. As such, crossover apertures 104may be formed from plenum 98 to cavity 102 even if crossover apertures104 are not formed at the exact intended location, i.e., crossoverapertures 104 do not have to be held to as tight of a tolerance ifcavity 102 provides a larger area in which crossover aperture 104 canjoin cavity 102. Likewise, in some embodiments, cavity 102 may beformed, shaped, and/or oriented to provide a sufficient target forforming ejection apertures 106 from outer surface 85 to cavity 102. Instill other embodiments, chambers 108 may be similarly formed, shaped,and/or oriented to provide a large target for forming crossoverapertures 104 to connect chamber 108 and plenum 98 and for formingejection apertures 106 to provide a passage from chamber 108 to outersurface 85 of airfoil 80. By positioning, sizing, and shaping cavity 102and/or chambers 108 as described, any tolerances required for drilling,machining, or otherwise forming crossover apertures 104 and ejectionapertures 106 may be accommodated without increasing the thickness oftrailing edge portion 91. Accordingly, the weight of airfoil 80 and thematerial required to produce airfoil 80 do not have to be increased toaccommodate tolerances in forming cooling passages 100 in trailing edgeportion 91 of airfoil 80. Moreover, optimal cooling passage size, shape,and/or position may decrease a time required to manufacture airfoil 80,e.g., by reducing the time required to drill, machine, or otherwise formcrossover apertures 104 and ejection apertures 106. Reducing thecomplexity and length of the manufacturing process also may decreasemanufacturing costs.

Although cooling passages 100 may be particularly beneficial along oradjacent trailing edge 90 of airfoil 80, cooling passages 100 may besuitable for any location on airfoil 80 and, for example, may be definedover pressure and suction sides 82, 84 of airfoil 80. In someembodiments, cavity 102, crossover apertures 104, and ejection apertures106 may all defined in either first portion 94 a or second portion 94 bof filler pack 94. In the embodiment illustrated in FIG. 5, cavity 102,crossover aperture 104, and ejection aperture 106 are each defined infirst portion 94 a of filler pack 94. Likewise, in embodiments such asthe embodiment illustrated in FIG. 7, chamber 108, crossover aperture104, and ejection aperture 106 may each be defined in one portion offiller pack 94. Thus, in some embodiments cooling passages 100, whetherdefined by cavity 102, crossover aperture 104, and ejection aperture 106or chamber 108, crossover aperture 104, and ejection aperture 106, maybe defined in only one portion of filler pack 94, e.g., in either firstportion 94 a or second portion 94 b of filler pack 94.

By defining cooling apertures 100 in only one half of filler pack 94, analignment requirement with respect to filler pack halves 94 a, 94 b canbe eliminated or avoided. That is, when the cooling passages 100 aredefined in only one portion of filler pack 94, the multiple portions offiller pack 94 do not have to be aligned to form cooling passages 100.Eliminating or avoiding a requirement to particularly align filler packportions can help simplify manufacturing and assembly of airfoil 80,e.g., by minimizing the machining or fabrication required to produce thevoids forming cooling passages 100, as well as by minimizing the stepsor processes to align filler pack portions.

Additionally or alternatively, by defining cooling apertures 100 in onlyone portion of filler pack 94, ejection aperture 106 may be biased toone side or the other. For example, as shown in FIG. 5, when ejectionaperture 106 is defined in filler pack first portion 94 a, whereejection aperture 106 exits at outer surface 85 may be biased topressure side 82 of airfoil 80, i.e., ejection aperture 106 may definean outlet 110 at pressure side 82. Similarly, when ejection aperture 106is defined in filler pack second portion 94 b, where ejection aperture106 exits at outer surface 85 may be biased to suction side 84 ofairfoil 80, i.e., ejection aperture 106 may define an outlet 110 atsuction side 84. Moreover, for airfoils 80 comprising a plurality ofcooling passages 100, outlets 110 may be defined at various axiallocations along pressure side 82 or suction side 84, i.e., outlets 110may not be radially aligned. As such, outlets 110 of cooling passages100 may be spaced apart generally along the axial direction A as well asthe radial direction R. Further, the use of multiple cooling passages100 at multiple locations of airfoil 80 may help enhance the surfacecooling provided by the cooling fluid flowing from each passage 100.

Various methods, techniques, and/or processes may be used to form cavity102, crossover apertures 104, ejection apertures 106, and chambers 108in airfoil 80. For example, in some embodiments, the portion ofcrossover aperture 104 defined through plenum plies 96 may be defined bycutting each individual plenum ply 96 before plenum plies 96 are laid upto form airfoil 80. In one embodiment, plies 96 are cut using aprecision Gerber cutter by Gerber Technology of Tolland, Connecticut. Inother embodiments, another type of cutter or other means for definingcut-outs in plies 96 may be used. As another example of forming voids inairfoil 80, crossover apertures 104 may be defined in plenum plies 96and filler pack 94 using electrical discharge machining (EDM), i.e., EDMdrilling.

In a further example, voids 102, 104, 106, and/or 108 may be formedusing one or more fugitive material inserts. That is, an insert madefrom a fugitive material may be in a desired form (e.g., shape, size,etc.) to define the corresponding void, e.g., cavity 102, crossoverapertures 104, ejection apertures 106, and/or chamber 108. The fugitivematerial insert is positioned within the lay-up as plenum plies 96,filler pack 94, and airfoil plies 92 are laid up to form airfoil 80. Insome embodiments, the insert may be formed of SiC fibers in a silicacarbide matrix. The insert may be one of various forms, such as a tapecast, a preformed silicon dioxide tube, or a rapid prototype polymercoating with boron nitride, and the insert may be formed in variousmanners, e.g., sprayed, screen printed, or injection molded. Forexample, the fugitive material insert may be a fugitive materialparticulate bound by polymer in a flexible tape. It may be desirablethat the fugitive material insert be a low melting metal or alloy thatmay melt during a burnout pyrolysis operation or melt infiltration of aCMC layup preform, to thereby leave a void in the preform. Inalternative embodiments, the fugitive material insert may be formed of ahigh temperature material that will not melt during the burnoutpyrolysis operation. For example, such fugitive materials include, butare not limited to, boron nitride (BN), silicon oxide, silicon oxidecoated with boron nitride, rare earth elements, rare earth elementscoated with boron nitride, rare earth oxides, rare earth oxides coatedwith boron nitride, rare earth silicate, rare earth silicate coated withboron nitride, elemental molybdenum, elemental molybdenum coated withboron nitride, molybdenum silicides, molybdenum silicides coated withboron nitride, gallium oxide, gallium nitride, indium oxide, indiumnitride, tin oxide, tin nitride, indium tin oxide (ITO), alkaline earthsilicates where the alkaline earth is magnesium, calcium, strontium,barium, and combinations thereof, alkaline earth aluminates, diamondpowder, diamond powder coated with boron nitride, or boron nitridecoated with carbon and mixtures and combinations thereof. All of thesehigh temperature materials may be placed into the CMC during layup as aflexible tape filled with powders of the high temperature materials.Alternately, all of these high temperature materials may also be placedinto the CMC during layup as a dense, flexible wire or an inflexible rodor tube. Such high temperature materials, after the CMC component ismelt infiltrated, may require a subsequent air heat treatment to oxidizethe high temperature material, a vacuum heat treatment, an inert gasheat treatment, an acid treatment, a base treatment, combinationsthereof, or alternating combinations thereof, to remove the fugitivematerial. Thus, the fugitive material may be removed by melting,dissolution, sublimation, evaporation, or the like.

Accordingly, various materials are suitable for use as the insert, suchas materials that exhibit non-wetting of the CMC preform, low or noreactivity with the constituents of the CMC preform, and/or arecompletely fusible and drainable at a temperature of a thermal treatmentperformed on the CMC preform. In one example embodiment, fugitivematerial inserts for defining ejection apertures 106 are formed of fusedsilicon dioxide (SiO₂) in a tubular shape. The tubes have, as anon-limiting example, an inner diameter of 10 mils and an outer diameterof 30 mils. The tubes may be positioned in an array within trailing edgeportion 91 of a layup of plies 92, 96 and filler pack(s) 94 for formingairfoil 80. Following a melt infiltration process, the fused silicondioxide is reduced to SiO. Such an insert will not wet or react with theconstituents of the preform. Additionally, the insert may melt and beallowed to drain from the preform during burnout, leaving the CMCpreform with voids forming ejection apertures 106.

FIG. 8 provides a chart illustrating an exemplary method 800 forfabricating airfoil 80. As represented at 802 in FIG. 8, plies 92, 96and filler pack(s) 94 are laid up in the form of airfoil 80, i.e., laidup in a desired shape to produce an airfoil preform assembly. The layupstep or portion of the process thus may be referred to as the layuppreforming step. The layup preforming step may comprise layeringmultiple plies or structures, such as plies pre-impregnated (pre-preg)with matrix material, pre-preg tapes, or the like, to form a desiredshape of the resultant CMC component, e.g., airfoil 80. The layers arestacked to form a layup or preform, which is a precursor to the CMCcomponent.

In some embodiments, multiple layups or preforms may be laid up togetherto form a preform assembly. More particularly, the layup portion ofmethod 800 depicted at 802 may include laying up multiple preformsand/or plies in an airfoil preform assembly 80P. Referring to FIG. 9, inan exemplary embodiment, the layup preforming step 802 may includeforming a plenum preform 96P and a filler pack preform 94P, which arelaid up with airfoil plies 92 and a second filler pack portion 94 b toproduce airfoil preform assembly 80P. More specifically, as shown at 902in FIG. 9, plenum plies 96 are laid up, e.g., in or on a layup tool,mandrel, or mold, to define a plenum preform 96P, illustrated in FIG.10. As shown in FIG. 10, plenum preform 96P generally defines the shapeof plenum 98 of airfoil 80. The plenum preform 96P may be compacted asillustrated at 904, and then processed in an autoclave as shown at 906.The compaction may be performed at atmosphere, i.e., at roomtemperature. The autoclave processing may be performed at reducedtemperature compared to a standard autoclave cycle such that plenumpreform 96P retains some flexibility and malleability after autoclaving.Such flexibility and malleability may help in laying up plenum preform96P with other preforms and plies to produce preform assembly 80P. Insome embodiments, the compaction and/or autoclaving steps 904, 906 maybe omitted, i.e., the compaction and autoclaving indicated at 904 and906 are optional, such that defining the plenum preform 96P compriseslaying up plenum plies 96 without additional processing. Further, inother embodiments, prior to, or as part of, laying up plenum plies 96 at902, plenum plies 96 may be cut to define at least a portion ofcrossover aperture(s) 104.

The layup preforming shown at 802 in FIG. 8 further may include forminga filler pack preform 94P, as shown in FIG. 9. As indicated at 908,filler pack material 94 is laid up, e.g., in or on a layup tool,mandrel, or mold, to define the filler pack preform 94P. Next, at 910,filler pack preform 94P is compacted, e.g., at atmosphere as describedabove with respect to the plenum preform. Then, as shown at 912, thefiller pack preform 94P is processed in an autoclave, e.g., at a reducedtemperature relative to a standard autoclave cycle such that filler packpreform 94P retains some flexibility and malleability after autoclaving.The flexibility and malleability may help in defining voids in thefiller pack preform as illustrated at 914 in FIG. 9. More particularly,after autoclaving, filler pack preform 94P is in a green state, andafter autoclaving at a reduced temperature, the green state filler packpreform 94P retains some flexibility and malleability that can assist infurther manipulation of the preform. For example, the voids formingcavity 102 and ejection aperture(s) 106 may be machined in the greenstate filler pack preform 94P, such that, in exemplary embodiments likethe embodiment of FIG. 5, filler pack preform 94P generally definesfiller pack first portion 94 a of airfoil 80. In other embodiments, thevoids forming chamber(s) 108 and ejection aperture(s) 106 may bemachined in green state filler pack preform 94P. The malleability ofgreen state preform 94P may help in forming voids 102, 106 or voids 108,106. In various embodiments, the voids may be formed using one or moreof laser drilling, EDM, cutting, or other machining methods. In otherembodiments, one or more of voids 102, 106, 108 may be formed usingfugitive material inserts and one or more of the processes or stepspreviously described.

Referring still to FIG. 9, as shown at 916, laying up the CMC materialto produce airfoil preform assembly 80P also may include preparingairfoil plies 92 for laying up with plenum preform 96P and filler packpreform 94P. It will be appreciated that, when laid up with filler packpreform 94P and plenum preform 96P, airfoil plies 92 generally define amajority of the shape of pressure and suction sides 82, 84 of theresultant airfoil 80 as described above.

Further, as shown at 918 in FIG. 9, laying up the CMC material to formairfoil preform assembly 80P may include laying up one or moreadditional filler pack(s). For example, referring to FIG. 10, a secondfiller pack portion 94 b and additional filler pack material 94 may belaid up with filler pack preform 94P and plenum preform 96P to furtherdefine trailing edge portion 91 and voids 102, 106 (or voids 108, 106 inembodiments utilizing chamber(s) 108 rather than cavity 102) withinairfoil 80 and to fill in any gaps between plenum preform 96P andairfoil plies 92. In particular embodiments, second filler pack portion94 b may be configured to fully define the shape of voids 102, 106 (orvoids 108, 106), i.e., the voids may be partially defined by filler packpreform 94P and partially defined by second filler pack portion 94 b.However, in some embodiments, second filler pack portion 94 b may beomitted as described above, such that trailing edge portion 91 ofairfoil 80 and cavity 102 and ejection aperture(s) 106 (or chamber(s)108 and ejection aperture(s) 106) within the trailing edge portion arefully defined by filler pack preform 94P. Further, as and if needed,additional filler pack(s) 94 may be positioned between airfoil plies 92and plenum preform 96P as shown in FIG. 10.

Accordingly, at layup preforming step 802 of method 800, the plenumpreform 96P, filler pack preform 94P, airfoil plies 92, and additionalfiller pack portions 94, 94 b may be laid up together to form airfoilpreform assembly 80P. In some embodiments, the layup preforming step 802also may comprise positioning one or more fugitive material insertswithin the layers to form one or more of voids 102, 106, 108 withinairfoil 80 as described above.

Next, airfoil preform assembly 80P is processed as shown at 804 in FIG.8. For example, airfoil preform assembly 80P may be processed in anautoclave using a standard autoclave process. As such, airfoil preformassembly 80P may be autoclaved at a higher temperature than filler packpreform 94P and plenum preform 96P as described above. After processing,if crossover aperture(s) 104 are not formed by cutouts in plenum plies96 as previously described, these apertures may be defined in the greenstate airfoil preform assembly 80P. For example, crossover aperture(s)104 may be EDM drilled from plenum 98 into the airfoil preform assembly,e.g., through plenum preform 96P and filler pack preform 94P, to theextent crossover aperture(s) 104 extend through filler pack 94. Invarious embodiments, the voids may be formed using one or more of laserdrilling, EDM, cutting, or other machining methods or using fugitivematerial inserts as previously described.

Next, as shown at 806 in FIG. 8, the airfoil preform assembly mayundergo a burn-out cycle, i.e., a burn-out cycle may be performed. In anexample burn-out cycle, any mandrel-forming materials, as well ascertain fugitive materials or other meltable materials such asadditional binders, are melted to remove such materials. Duringburn-out, the CMC airfoil preform assembly may be positioned to allowthe melted materials to run out of the preform and thus remove thematerials from the preform.

Then, as illustrated at 808, the CMC airfoil preform assembly may besubjected to one or more post-processing cycles for densification of thepreform assembly. Densification may be performed using any knowndensification technique including, but not limited to, Silcomp, meltinfiltration (MI), chemical vapor infiltration (CVI), polymerinfiltration and pyrolysis (PIP), and oxide/oxide processes.Densification can be conducted in a vacuum furnace having an establishedatmosphere at temperatures above 1200° C. to allow silicon or othermaterials to melt-infiltrate into the preform component.

Additionally or alternatively, after burn-out and post-processing steps806, 808, airfoil 80 may be manipulated mechanically or chemically asshown at 810 in FIG. 8 to remove any remaining fugitive materialinserted into the preformed shape during layup preforming step 802. Insome cases, the heat treatment may be used to oxidize the insert to anoxide that may be melted or dissolved in an acid or base. In otherembodiments, the insert may be directly dissolved in acid or base, orotherwise chemically dissolved. In further embodiments, the insert maybe sublimed or evaporated in a vacuum heat treatment. In still otherembodiments, the insert may be oxidized and subsequently sublimed orevaporated in a vacuum heat treatment. Mechanical methods may be used tomechanically remove the insert, and such mechanical methods may or maynot be used with any of the previously described methods. Variouschemical methods may be utilized as well.

After any remaining fugitive material is removed at step 810, airfoil 80may be finish machined as shown at 812. Finish machining may includeclearing the formed features, such as crossover apertures 104 andejection apertures 106, with wire to, e.g., ensure proper flow throughthe voids. Subsequently, an environmental barrier coating (EBC) may beapplied to the airfoil 80, as shown at 814. Prior to applying thecoating, a stop-off comb may be inserted into ejection apertures 106 toprevent the coating from blocking the cooling passages 100.

Method 800 is provided by way of example only; it will be appreciatedthat some steps or portions of method 800 may be performed in anotherorder. Additionally, other methods of fabricating or forming airfoil 80may be used as well. In particular, other processing cycles, e.g.,utilizing other known methods or techniques for compacting CMC plies,may be used. Further, airfoil 80 may be post-processed using a meltinfiltration process, a chemical vapor infiltration process, a matrix ofpre-ceramic polymer fired to obtain a ceramic matrix, or anycombinations of these or other known processes.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal language of the claims.

1.-15. (canceled)
 16. A method for forming an airfoil for a gas turbineengine, the method comprising: laying up a ceramic matrix compositematerial to form an airfoil preform assembly, the airfoil preformassembly including opposite pressure and suction sides extendingradially along a span, opposite leading and trailing edges extendingradially along the span, the pressure and suction sides extendingaxially between the leading and trailing edges, and a plenum definedwithin the airfoil preform assembly; and processing the airfoil preformassembly to produce the airfoil, wherein a cooling passage is definedwithin the airfoil, the cooling passage defined from the plenum to thetrailing edge of the airfoil.
 17. The method of claim 16, wherein thecooling passage comprises a crossover aperture defined from the plenumto a cavity defined by a plenum preform, and wherein the cooling passagefurther comprises an ejection aperture defined within a filler packpreform from the trailing edge to the cavity.
 18. The method of claim16, further comprising: performing a burn-out cycle after processing theairfoil preform assembly.
 19. The method of claim 16, furthercomprising: post-processing the airfoil preform assembly afterprocessing the airfoil preform assembly.
 20. The method of claim 19,further comprising: finish machining the airfoil preform assembly afterpost-processing the airfoil preform assembly.
 21. The method of claim19, wherein post-processing the airfoil preform assembly densifies theairfoil preform assembly.
 22. The method of claim 16, wherein laying upthe ceramic matrix composite material to form the airfoil preformassembly includes laying up a fugitive material insert within layers ofthe ceramic matrix composite material to define a void in the airfoil.23. The method of claim 22, further comprising, after processing theairfoil preform assembly: mechanically manipulating the airfoil toremove fugitive material of the fugitive material insert.
 24. The methodof claim 22, further comprising, after processing the airfoil preformassembly: chemically manipulating the airfoil to remove fugitivematerial of the fugitive material insert.
 25. The method of claim 16,wherein laying up the ceramic matrix composite material to form theairfoil preform assembly comprises forming a plenum preform and a fillerpack preform and laying up the plenum preform and the filler packpreform with a plurality of airfoil plies.
 26. The method of claim 25,wherein forming the plenum preform comprises laying up a plurality ofplenum plies to define the plenum preform, compacting the plenumpreform, and processing the plenum preform in an autoclave.
 27. Themethod of claim 25, wherein forming the filler pack preform compriseslaying up filler pack material to define the filler pack preform,compacting the filler pack preform, and processing the filler packpreform in an autoclave.
 28. The method of claim 27, further comprising:defining one or more voids in the filler pack preform prior to laying upthe filler pack preform with the plenum preform and the plurality ofairfoil plies.
 29. The method of claim 28, wherein the filler packpreform is machined to define the one or more voids are defined in thefiller pack preform.
 30. The method of claim 28, wherein one or morefugitive material inserts are laid up with the filler pack material todefine the one or more voids.
 31. A method for forming an airfoil for agas turbine engine, the method comprising: laying up a ceramic matrixcomposite material to form an airfoil preform assembly, the airfoilpreform assembly including opposite pressure and suction sides extendingradially along a span, the pressure and suction sides defining an outersurface of the airfoil, opposite leading and trailing edges extendingradially along the span, the pressure and suction sides extendingaxially between the leading and trailing edges, a plenum defined withinthe airfoil preform assembly, a plurality of airfoil plies defining inpart the pressure and suction sides, and a filler pack preform definingthe trailing edge, the filler pack preform having a forward portion andan aft portion, the forward portion positioned against the plenum andinward of the airfoil plies such that the airfoil plies define the outersurface of the airfoil at the forward portion, the aft portion definingthe outer surface of the airfoil at the trailing edge; and processingthe airfoil preform assembly to produce the airfoil, wherein a coolingpassage is defined within the airfoil, the cooling passage defined fromthe plenum to the trailing edge of the airfoil, at least a portion ofthe cooling passage defined in the filler pack preform.
 32. The methodof claim 31, wherein the cooling passage comprises a radially extendingcavity defined in the forward portion of the filler pack preform, acrossover aperture defined from the plenum to the cavity to facilitatethe flow of cooling fluid from the plenum to the cavity, and an ejectionaperture defined in the filler pack preform from the cavity to the outersurface of the airfoil adjacent the trailing edge, and wherein theejection aperture is in fluid communication with the cavity to directthe flow of cooling fluid from the plenum to the outer surface of theairfoil.
 33. The method of claim 31, wherein a plenum preform definesthe plenum, and wherein laying up the ceramic matrix composite materialto form the airfoil preform assembly comprises laying up the plenumpreform and the filler pack preform with the plurality of airfoil plies.34. The method of claim 31, further comprising, after processing theairfoil preform assembly: performing a burn-out cycle; densifying theairfoil preform assembly; and manipulating the airfoil preform assemblyto remove fugitive material used to define the cooling passage.
 35. Amethod for forming an airfoil for a gas turbine engine, the methodcomprising: laying up a ceramic matrix composite material to form anairfoil preform assembly comprising preparing a plurality of airfoilplies, forming a plenum preform defining a plenum, forming a filler packpreform, and preparing one or more filler packs; and processing theairfoil preform assembly to produce the airfoil, wherein the airfoilcomprises opposite pressure and suction sides extending radially along aspan and opposite leading and trailing edges extending radially alongthe span, the pressure and suction sides extending axially between theleading and trailing edges, and wherein a cooling passage is definedwithin the airfoil, the cooling passage defined from the plenum to thetrailing edge of the airfoil.